Outwardly-looking radar, infrared, and/or visible-light sensors built into vehicles such as aircraft or missiles are usually protected by a covering termed a dome. The dome serves as a window that transmits the radiation sensed by the sensor. The dome can also act as a structural element that protects the sensor and that can carry aerodynamic loadings. In many cases, the dome can protect a forward-looking sensor, wherein the dome bears large aerostructural loadings.
Where the vehicle moves relatively slowly, as in the case of helicopters, subsonic aircraft, and ground vehicles, some domes are made of nonmetallic organic materials which have good energy transmission and low signal distortion, and can support small-to-moderate structural loadings at low-to-intermediate temperatures. For those vehicles that fly much faster, such as hypersonic aircraft or missiles flying in the Mach 3-20 range, nonmetallic organic materials are inadequate for use in domes because aerodynamic friction heats the dome above the maximum operating temperature of the organic material.
In such cases, the dome is typically made of a ceramic material that can withstand elevated temperatures and that has good energy transmission characteristics. However, existing ceramics, such as sapphire, have the shortcoming that they are relatively brittle and non-elastic. The likelihood of fracture can be increased by the presence of small surface defects in the ceramic and externally imposed stresses and strains. The ceramic dome can be hermetically attached to the body of the missile, which is typically made of a metal with high-temperature strength, such as a titanium alloy.
Ceramic material has a relatively low coefficient of thermal expansion (CTE), while the metal missile body typically has a relatively high CTE. Changing the temperature of the missile body and dome can result in a CTE mismatch, which can create or induce strain between the dome and the missile body when the two are joined. This can greatly increase the propensity of the dome to fracture in a brittle manner and can lead to failure of the sensor and ultimately failure of the missile. In one typical example, the dome and the missile body are joined by brazing at approximately 1000 degrees F. At this temperature, there is effectively little to no strain in the joint due to a CTE mismatch. A temperature change can occur as the parts cool from the joining temperature. Additional temperature changes can occur, for example, when the missile is carried on board a launch aircraft or during service, in which the temperature can drop to −55 degrees F. The difference in temperature between 1000 degrees F. and −55 degrees F. can create the greatest CTE mismatch that the dome and missile body experience and, therefore, the greatest strain between the dome and the missile body. In other words, the maximum CTE mismatch stress occurs at low temperatures, when the substantially “zero stress state” at braze temperature is at its greatest difference.
To account for this CTE mismatch between the dome and missile body, some designs comprise multiple parts coupled by brazing and include transition elements to reduce the severity of CTE mismatching in stages. For example, a transition element may have an intermediate CTE relative to the dome and missile body to allow the dome to be coupled indirectly to the missile body. The result is a complex design that may also require additional aerodynamic components and sealing of joints and gaps between components, such as with polysulfide.
Reference will now be made to the exemplary embodiments illustrated, and specific language will be used herein to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended.